Gas turbine having a peripheral ring segment including a recirculation channel

Abstract

A gas turbine having at least one compressor, at least one combustion chamber, and at least one turbine, the or each compressor and/or the or each turbine having a rotor that includes rotor blades surrounded by a stationary housing, and a run-in coating being assigned to the housing is disclosed. The gas turbine includes at least one channel which is configured to apply a pressure prevailing on the high-pressure side of the blades of a rotor to a low-pressure side of the same in the area of a gap between the radially outer ends of the blades and the housing and thereby prevent a flow through the gap.

1. A gas turbine comprising:
 at least one compressor; 
 at least one combustion chamber; and 
 at least one turbine, at least one of the at least one compressor and the least one turbine comprising a rotor and a stationary housing, the rotor including rotor blades, the stationary housing including a run-in coating and at least one channel, the stationary housing surrounding the rotor blades such that there is a gap between radially outer ends of the rotor blades and the stationary housing, the stationary housing further including a peripheral ring segment forming a substrate for the run-in coating; 
 wherein the at least one channel is configured to apply a pressure prevailing on a high-pressure side of the rotor blades to a low-pressure side of the rotor blades in an area of the gap to prevent a flow through the gap, at least a portion of the at least one channel extending within the peripheral ring, the at least one channel opening on the low-pressure side into the gap in a first area of the run-in coating at a low-pressure side opening, and opening on the high-pressure side in a second area of the peripheral ring segment outside of the gap at a high-pressure side opening. 
 
     
     
       2. The gas turbine as recited in  claim 1  wherein the run-in coating is gas-permeable and has an open-cell structure. 
     
     
       3. The gas turbine as recited in  claim 1  wherein the run-in coating is a metal foam. 
     
     
       4. The gas turbine as recited in  claim 1  wherein a cross section of the at least one channel is dimensioned in such a way that air flowing through the at least one channel acts as sealing air in the area of the gap. 
     
     
       5. The gas turbine as recited in  claim 1  wherein the gas turbine is an aircraft engine. 
     
     
       6. The gas turbine as recited in  claim 1  wherein each blade of the blades has a radial outer edge extending axially between the low-pressure side and the high-pressure side. 
     
     
       7. The gas turbine as recited in  claim 6  wherein the low-pressure side opening is spaced axially from the outer edge. 
     
     
       8. The gas turbine as recited in  claim 6  wherein the high-pressure side opening is spaced axially from and not in contact with the run-in coating. 
     
     
       9. The gas turbine as recited in  claim 1  wherein the high-pressure side opening is spaced axially from and not in contact with the run-in coating. 
     
     
       10. The gas turbine as recited in  claim 1  wherein each blade of the blades has a radial outer edge extending axially between a low pressure side and a high pressure side and the run-in coating has a recess into which the radial outer edge extends. 
     
     
       11. The gas turbine as recited in  claim 10  wherein the recess has a first radially extending surface on the low-pressure side, a second radially-extending surface on the high-pressure side and an axially-extending surface between the first and second radially-extending surfaces, the low-pressure side opening opening into the first-radially extending surface. 
     
     
       12. The gas turbine as recited in  claim 1  wherein the channel at the low-pressure side opening extends toward the low-pressure side.